RAE 104 AIRFOIL (rae104-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: RAE 104 AIRFOIL (rae104-il) Reynolds number: 1,000,000 Max Cl/Cd: 58.13 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rae104-il-1000000.txt Download as CSV file: xf-rae104-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: RAE 104 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.000 -0.7922 0.06179 0.05998 -0.0378 1.0000 0.0070
-11.750 -0.8221 0.05450 0.05254 -0.0421 1.0000 0.0067
-11.500 -0.8529 0.04824 0.04606 -0.0432 1.0000 0.0068
-11.250 -0.8859 0.04314 0.04074 -0.0416 1.0000 0.0065
-11.000 -0.9094 0.03936 0.03671 -0.0380 1.0000 0.0066
-10.750 -0.9379 0.03562 0.03270 -0.0319 1.0000 0.0065
-10.500 -0.9511 0.03250 0.02928 -0.0271 1.0000 0.0067
-10.250 -0.9548 0.02977 0.02624 -0.0232 1.0000 0.0068
-10.000 -0.9513 0.02765 0.02387 -0.0200 1.0000 0.0069
-9.750 -0.9409 0.02633 0.02236 -0.0176 1.0000 0.0071
-9.500 -0.9276 0.02535 0.02123 -0.0155 1.0000 0.0072
-9.250 -0.9232 0.02270 0.01825 -0.0121 1.0000 0.0074
-9.000 -0.9169 0.02010 0.01535 -0.0088 1.0000 0.0076
-8.750 -0.9053 0.01854 0.01361 -0.0062 1.0000 0.0078
-8.500 -0.8911 0.01757 0.01255 -0.0039 1.0000 0.0081
-8.250 -0.8764 0.01683 0.01174 -0.0017 1.0000 0.0083
-8.000 -0.8624 0.01616 0.01101 0.0007 1.0000 0.0086
-7.750 -0.8325 0.01539 0.01015 -0.0003 0.9985 0.0090
-7.500 -0.8001 0.01475 0.00945 -0.0017 0.9962 0.0096
-7.250 -0.7681 0.01415 0.00878 -0.0031 0.9936 0.0103
-7.000 -0.7368 0.01370 0.00828 -0.0041 0.9901 0.0107
-6.750 -0.7087 0.01238 0.00681 -0.0048 0.9860 0.0115
-6.500 -0.6767 0.01164 0.00602 -0.0061 0.9833 0.0128
-6.250 -0.6470 0.01122 0.00557 -0.0069 0.9775 0.0139
-6.000 -0.6133 0.01083 0.00515 -0.0084 0.9734 0.0152
-5.750 -0.5797 0.01031 0.00457 -0.0100 0.9689 0.0168
-5.500 -0.5491 0.00982 0.00407 -0.0109 0.9601 0.0201
-5.250 -0.5167 0.00955 0.00376 -0.0121 0.9514 0.0225
-5.000 -0.4882 0.00913 0.00330 -0.0124 0.9396 0.0283
-4.750 -0.4613 0.00890 0.00303 -0.0123 0.9268 0.0341
-4.500 -0.4369 0.00861 0.00276 -0.0117 0.9138 0.0473
-4.250 -0.4131 0.00833 0.00252 -0.0110 0.9016 0.0677
-4.000 -0.3900 0.00800 0.00230 -0.0102 0.8905 0.1023
-3.750 -0.3681 0.00758 0.00208 -0.0092 0.8806 0.1599
-3.500 -0.3468 0.00712 0.00186 -0.0081 0.8714 0.2336
-3.250 -0.3269 0.00654 0.00163 -0.0069 0.8627 0.3347
-3.000 -0.3083 0.00590 0.00141 -0.0053 0.8550 0.4553
-2.750 -0.2905 0.00527 0.00123 -0.0035 0.8472 0.5794
-2.500 -0.2681 0.00505 0.00120 -0.0023 0.8411 0.6547
-2.250 -0.2423 0.00497 0.00117 -0.0019 0.8352 0.6832
-2.000 -0.2159 0.00495 0.00114 -0.0016 0.8297 0.7018
-1.750 -0.1892 0.00493 0.00113 -0.0013 0.8244 0.7167
-1.500 -0.1624 0.00491 0.00111 -0.0011 0.8190 0.7299
-1.250 -0.1354 0.00493 0.00109 -0.0009 0.8139 0.7411
-1.000 -0.1084 0.00490 0.00110 -0.0007 0.8084 0.7511
-0.750 -0.0813 0.00490 0.00108 -0.0005 0.8025 0.7598
-0.500 -0.0542 0.00492 0.00108 -0.0004 0.7966 0.7684
-0.250 -0.0271 0.00490 0.00108 -0.0002 0.7898 0.7758
0.000 0.0000 0.00493 0.00108 0.0000 0.7832 0.7831
0.250 0.0271 0.00490 0.00108 0.0002 0.7757 0.7898
0.500 0.0542 0.00492 0.00108 0.0004 0.7684 0.7965
0.750 0.0813 0.00490 0.00109 0.0005 0.7599 0.8025
1.000 0.1084 0.00490 0.00110 0.0007 0.7509 0.8084
1.250 0.1354 0.00493 0.00109 0.0009 0.7411 0.8139
1.500 0.1624 0.00491 0.00111 0.0011 0.7300 0.8190
1.750 0.1892 0.00493 0.00113 0.0013 0.7167 0.8244
2.000 0.2159 0.00495 0.00114 0.0016 0.7017 0.8297
2.250 0.2423 0.00497 0.00117 0.0019 0.6834 0.8352
2.500 0.2681 0.00505 0.00120 0.0023 0.6542 0.8411
2.750 0.2906 0.00527 0.00123 0.0035 0.5804 0.8472
3.000 0.3083 0.00590 0.00141 0.0053 0.4550 0.8550
3.250 0.3268 0.00655 0.00163 0.0069 0.3336 0.8627
3.500 0.3468 0.00712 0.00186 0.0081 0.2339 0.8714
3.750 0.3681 0.00758 0.00208 0.0092 0.1601 0.8806
4.000 0.3901 0.00800 0.00230 0.0102 0.1028 0.8905
4.250 0.4131 0.00833 0.00252 0.0110 0.0676 0.9016
4.500 0.4368 0.00862 0.00276 0.0117 0.0469 0.9138
4.750 0.4614 0.00889 0.00302 0.0123 0.0345 0.9269
5.000 0.4882 0.00913 0.00330 0.0124 0.0285 0.9397
5.250 0.5166 0.00957 0.00377 0.0121 0.0226 0.9515
5.500 0.5491 0.00981 0.00404 0.0109 0.0202 0.9602
5.750 0.5799 0.01027 0.00452 0.0099 0.0168 0.9687
6.000 0.6133 0.01081 0.00513 0.0084 0.0152 0.9733
6.250 0.6469 0.01122 0.00558 0.0069 0.0140 0.9775
6.500 0.6766 0.01164 0.00602 0.0062 0.0128 0.9834
6.750 0.7086 0.01237 0.00681 0.0048 0.0115 0.9860
7.000 0.7368 0.01367 0.00825 0.0041 0.0107 0.9901
7.250 0.7679 0.01417 0.00881 0.0031 0.0102 0.9936
7.500 0.8000 0.01476 0.00946 0.0018 0.0096 0.9961
7.750 0.8324 0.01540 0.01016 0.0003 0.0089 0.9984
8.000 0.8623 0.01619 0.01105 -0.0007 0.0086 1.0000
8.250 0.8764 0.01685 0.01176 0.0017 0.0083 1.0000
8.500 0.8912 0.01759 0.01257 0.0039 0.0081 1.0000
8.750 0.9051 0.01861 0.01370 0.0063 0.0079 1.0000
9.000 0.9165 0.02019 0.01545 0.0089 0.0076 1.0000
9.250 0.9253 0.02233 0.01783 0.0118 0.0075 1.0000
9.500 0.9220 0.02631 0.02228 0.0162 0.0073 1.0000
9.750 0.9348 0.02727 0.02339 0.0183 0.0072 1.0000
10.000 0.9489 0.02799 0.02424 0.0203 0.0070 1.0000
10.250 0.9456 0.03092 0.02749 0.0242 0.0070 1.0000
10.500 0.9491 0.03272 0.02952 0.0273 0.0068 1.0000
10.750 0.9367 0.03572 0.03280 0.0321 0.0066 1.0000
11.000 0.9133 0.03894 0.03627 0.0376 0.0067 1.0000
11.250 0.8892 0.04276 0.04033 0.0414 0.0066 1.0000
11.500 0.8534 0.04842 0.04627 0.0432 0.0066 1.0000
11.750 0.8247 0.05431 0.05236 0.0421 0.0066 1.0000
12.000 0.7928 0.06212 0.06034 0.0376 0.0068 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RAE 104 AIRFOIL (rae104-il)